Methods and system for recuperated circumferential cooling of integral turbine nozzle and shroud assemblies

ABSTRACT

A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH & DEVELOPMENT

The U.S. Government may have certain rights in this invention pursuantto contract number N00019-04-C-0093.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines and, moreparticularly, to methods and systems for cooling integral turbine nozzleand shroud assemblies.

One known approach to increase the efficiency of gas turbine enginesrequires raising the turbine operating temperature. However, asoperating temperatures are increased, the thermal limits of certainengine components may be exceeded, resulting in reduced service lifeand/or material failure. Moreover, the increased thermal expansion andcontraction of components may adversely affect component clearancesand/or component interfitting relationships. Consequently, conventionalcooling systems have been incorporated into gas turbine engines tofacilitate cooling such components to avoid potentially damagingconsequences when exposed to elevated operating temperatures.

It is known to extract, from the main airstream, air from the compressorfor cooling purposes. To facilitate maintaining engine operatingefficiency, the volume of cooling air extracted is typically limited toa small percentage of the total main airstream. As such, this requiresthat the cooling air be utilized with the utmost efficiency in order tofacilitate maintaining the temperatures of components within safelimits.

For example, one component that is subjected to high temperatures is theshroud assembly located immediately downstream of the high pressureturbine nozzle extending from the combustor. The shroud assembly extendscircumferentially about the rotor of the high pressure turbine and thusdefines a portion of the outer boundary (flow path) of the main gasstream flowing through the high pressure turbine. Gas turbine engineefficiency may be negatively affected by a fluctuation in turbine bladetip clearance measured between a radially outer surface of the turbineblade and a radially inner surface of the shroud assembly. Duringtransient engine operation, turbine blade tip clearance is a function ofa difference in radial displacement of the turbine rotor blade and theshroud assembly. The turbine rotor typically has a larger mass than thestationary shroud system and, thus, during turbine operation, theturbine rotor typically has a slower thermal response than the shroudassembly. When the difference in the rotor blade radial displacement andthe shroud assembly radial displacement is too great, the bladeclearance is increased, which may result in reducing engine efficiency.

Moreover, during engine operation, a gap may be defined between atrailing edge of the high pressure turbine nozzle outer band and aleading edge of the adjacent shroud segment. Cooling air, including,without limitation, nozzle leakage and/or purge flow, enters the gap andflows into the main gas stream channeled through the high pressureturbine. More specifically, because known nozzle outer band trailingedges and shroud leading edges have a simple 90° corner, the gap opensdirectly into the main gas stream. During engine operation, as the maingas stream flows through the nozzle vanes, a circumferential gaspressure variation may be created downstream from the vane trailingedge. This circumferential gas pressure variation may cause localizedhot gas ingestion into the gap between the outer band and the shroudsegment. As a result, cooling air flowing through the gap may noteffectively cool the downstream shroud segment.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method for cooling a shroud segment of a gas turbineengine is provided. The method includes providing a turbine shroudassembly including a shroud segment having an inner surface and aleading edge that is substantially perpendicular to the inner surface,and coupling a turbine nozzle to the turbine shroud segment such that agap is defined between an aft edge of an outer band of the turbinenozzle and the leading edge. The method also includes directing coolingair into the gap, circumferentially mixing the cooling air in a plenumdefined within the leading edge to substantially uniformly distributethe cooling air throughout the gap, and directing the cooling air in thegap through at least one cooling hole formed between the plenum and theinner surface.

In a further aspect, a turbine nozzle and shroud assembly for a gasturbine engine is provided. The assembly includes a shroud segmentincluding a leading edge and an inner surface that is substantiallyperpendicular to the leading edge, and a turbine nozzle including anouter band including an aft edge. The turbine nozzle is upstream fromthe shroud segment and is coupled with the shroud segment such that agap is defined between the aft edge and the leading edge. The shroudsegment includes a circumferential plenum defined within the leadingedge. The plenum is configured to substantially uniformly distributecooling air throughout the gap. The shroud segment also includes atleast one cooling hole formed between the plenum and the inner surface.The at least one cooling hole is configured to direct the cooling air inthe gap toward the hot gas flow path.

In another aspect, a cooling system for a gas turbine engine isprovided. The gas turbine engine includes a shroud segment having aleading edge and an inner surface that is substantially perpendicular tothe leading edge, and a turbine nozzle that is upstream from the shroudsegment and that includes an outer band having an aft edge. The coolingsystem includes a circumferential plenum defined within the leadingedge. The plenum is configured to substantially uniformly distributecooling air throughout a gap defined between the aft edge and theleading edge. The cooling system also includes at least one cooling holeformed between the plenum and the inner surface. The at least onecooling hole is configured to direct the cooling air from the gaptowards a hot gas flow path defined within the gas turbine engine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side view of an exemplary shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.

FIG. 2 is a side view of an alternative shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.

FIG. 3 is a schematic view of a portion of an exemplary turbine nozzleand shroud assembly.

FIG. 4 is a schematic view of an alternative embodiment of a portion ofa turbine nozzle and shroud assembly.

DETAILED DESCRIPTION OF THE INVENTION

The present invention provides a turbine shroud cooling system for filmcooling a shroud segment. The turbine shroud cooling system facilitatesforming a barrier between the hot gas flow path flowing through the highpressure turbine and cooling air flowing through a gap defined betweenthe turbine nozzle and the shroud segment. More specifically, anextended lip at a trailing edge of the outer band facilitates formingthe barrier between the hot gas flow path and the gap defined between anouter band of the turbine nozzle and the shroud segment. Further, theextended lip facilitates pressurizing the gap to facilitate preventingor limiting hot gas injection into the gap. In one embodiment, theextended lip forms an axial aft facing film cooling slot in parallelwith a rounded corner portion of the shroud leading edge to facilitatefilm cooling the downstream shroud segment.

Although the present invention is described below in reference to itsapplication in connection with cooling a shroud assembly of an aircraftgas turbine, it should be apparent to those skilled in the art andguided by the teachings herein provided that with appropriatemodification, the cooling system or assembly of the present inventioncan also be suitable to facilitate cooling other turbine enginecomponents, such as, but not limited to, the nozzle and/or vanesections.

FIG. 1 is a side view of an exemplary shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.FIG. 2 is a side view of an alternative shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.To facilitate controlling shroud assembly thermal response and/or shroudassembly displacement during transient engine operation, in theexemplary embodiment, a turbine engine cooling assembly 108 includes ashroud assembly, generally indicated as 110, for a high pressure turbinesection 112 and a low pressure turbine section 114 of a gas turbineengine. It should be apparent to those skilled in the art and guided bythe teachings herein provided that turbine engine cooling assembly 108may be suitable to facilitate cooling other sections of the gas turbineengine, such as, but not limited to, a nozzle section and/or a vanesection.

Shroud assembly 110 includes turbine engine cooling components in theform of shroud segments 130. Each shroud segment 130 includes a forwardmounting hook 132 at a circumferential leading edge 133 of shroudsegment 130. Shroud segment 130 also includes a midsection mounting hook134 and an aft mounting hook 136 adjacent to a circumferential trailingedge 137 of shroud segment 130.

A plurality of shroud segments 130 are arranged circumferentially in agenerally known fashion to form an annular segmented shroud. Shroudsegments 130 define an annular clearance between high pressure turbineblades (not shown) and a radially inner surface 138 of a high pressureturbine section of shroud segments 130, and between low pressure turbineblades (not shown) and a radially inner surface 140 of a low pressureturbine section of shroud segment 130. A plurality of segmented shroudsupports 144 interconnect shroud segments 130. Each shroud support 144circumferentially spans and supports adjacent shroud segments 130. Inalternative embodiments, shroud supports 144 are modified to support anysuitable number of shroud segments 130 less than or greater than twoshroud segments 130. In the exemplary embodiment, shroud assembly 110includes twenty-six (26) shroud segments 130 and thirteen (13) shroudsupports 144, although any suitable number of shroud segments 130 and/orshroud supports 144 may be utilized in alternative embodiments.

Each shroud support 144 includes a forward section 146, a midsection 148and an aft section 150 that form respective forwardly projecting hangers152, 154 and 156. Mounting hooks 132, 134 and 136 are received bycooperating hangers 152, 154 and 156, respectively, in tongue-in-groove,or hook-in-hanger, interconnections such that shroud support 144supports respective shroud segments 130.

Shroud assembly 110 includes an annular shroud ring structure 158 thatsupports shroud supports 144. In one embodiment, shroud ring structure158 is a one-piece, continuous annular shroud ring structure. A radialposition of each shroud support 144, as well as of each shroud segment130, is closely controlled by only two annular position control rings162 and 164 formed on shroud ring structure 158. In contrast toconventional shroud ring structures, to facilitate reducing or limitinga weight of shroud assembly 110, shroud ring structure 158 includes onlytwo position control rings 162 and 164. A midsection position controlring 162 includes an axially forwardly projecting hanger 166 thatreceives and/or cooperates with a rearwardly projecting mounting hook167 formed by support structure midsection 148 in a firstcircumferential tongue-in-groove or hook-in-hanger interconnection. Anaft position control ring 164 includes an axially forwardly projectinghanger 168 that receives and/or cooperates with a rearwardly projectingmounting hook 169 of support structure aft section 150 in secondcircumferential tongue-in-groove or hook-in-hanger interconnection.

In the exemplary embodiment, hangers 166 and/or 168 are in direct axialalignment, i.e., aligned generally in the same radial plane, withrespective hanger 154 and hanger 156 to facilitate maximizing the radialsupport and/or radial position control provided to shroud support 144and, thus, corresponding shroud segments 130. This alignment orientationfacilitates increasing the rigidity of the entire shroud supportassembly. In an alternative embodiment, shown in FIG. 2, hanger 166and/or hanger 168 are in an offset axial alignment, i.e., not alignedgenerally in the same radial plane, with respective hanger 154 andhanger 156. In the exemplary embodiment, shroud ring structure 158 isbolted to the combustor case (not shown) at an aft end of shroud ringstructure 158. Shroud ring structure 158 is cantilevered away fromleading edge 133 at the combustor case interface. As such, midsectionposition control ring 162 is positioned several inches away from thecombustor aft flange (not shown), and is thereby divorced from anynon-uniform circumferential variations in radial deflection in thecombustor case.

In the exemplary embodiment, high pressure cooling air 170 is extractedfrom a compressor (not shown) positioned upstream of shroud assembly110. A first portion 171 of high pressure cooling air 170 extracted fromthe compressor facilitates cooling high pressure turbine section 112. Asecond portion 172 of high pressure cooling air 170 extracted from thecompressor facilitates cooling low pressure turbine section 114.Referring further to FIG. 1, directional arrows corresponding to firstportion 171 and second portion 172 illustrate at least a portion of aflow path of first portion 171 of high pressure cooling air 170 througha high pressure turbine section active convection cooling zone 173 andsecond portion 172 of high pressure cooling air 170 through a lowpressure turbine section active convection cooling zone 186 (describedbelow), respectively.

In this embodiment, first portion 171 of high pressure cooling air 170is metered into a first or high pressure turbine section activeconvection cooling zone 173. More specifically, first portion 171 ofhigh pressure cooling air 170 is metered through at least one highpressure turbine section (HPTS) feed hole 174 defined in shroud support144. First portion 171 of high pressure cooling air 170 impinges againsta pan-shaped HPTS impingement baffle 175 positioned within high pressureturbine section active convection cooling zone 173. Baffle 175 iscoupled to shroud support 144 and thus at least partially defines anupper HPTS cavity or plenum 176. First portion 171 of high pressurecooling air 170 is then metered through a plurality of perforations 177formed in impingement baffle 175 as cooling air into a lower HPTS cavityor plenum 178 defined in shroud segment 130, wherein the cooling airimpinges against a backside 179 of shroud segment 130. A portion, suchas spent impingement cooling air 180, of high pressure cooling air exitsplenum 178 through a plurality of forwardly directed cooling openings181 defined at, or near, shroud segment leading edge 133 configured tofacilitate purging a gap 182 defined between high pressure turbinenozzle outer band 183 and leading edge 133. A portion 184 of highpressure cooling air is metered through a plurality of rearwardlydirected cooling openings 185 defined in shroud segment 130 tofacilitate film cooling inner surface 138 and/or 140. Spent impingementcooling air 180 of high pressure cooling air exiting cooling openings181 facilitates preventing or limiting hot gas injection orrecirculation into shroud assembly 110 at leading edge 133.

Second portion 172 of high pressure cooling air 170 extracted from thecompressor facilitates cooling low pressure turbine section 114. In thisembodiment, second portion 172 of high pressure cooling air 170 ismetered into a second or low pressure turbine section active convectioncooling zone 186. More specifically, second portion 172 of high pressurecooling air 170 is metered through at least one low pressure turbinefeed hole 187 defined in shroud support 144. Second portion 172 of highpressure cooling air 170 impinges against a pan-shaped low pressureturbine section (LPTS) impingement baffle 188 positioned within lowpressure turbine section active convection cooling zone 186. Baffle 188is coupled to shroud support 144, and thus at least partially defines anupper LPTS cavity or plenum 189. Second portion 172 of high pressurecooling air 170 is then metered through perforations 190 defined inimpingement baffle 188 and into a lower LPTS cavity or plenum 191wherein high pressure cooling air impinges against a backside 192 ofshroud segment 130. Cooling air 193 exits plenum 191 through a pluralityof rearwardly directed cooling openings 194 defined through shroudsegment 130, to facilitate film cooling radially inner surface 140 oftrailing edge 137 of shroud segment 130 downstream.

As shown in FIG. 1, high pressure cooling air 170 is initially directedinto a duct 204 defined at least partially between high pressure turbinenozzle outer band 183 and the portion of shroud ring structure 158forming midsection position control ring 162. High pressure cooling air170 is separated within duct 204 into first portion 171, and into secondportion 172, as high pressure cooling air 170 is directed through duct204. First portion 171 of high pressure cooling air 170 is meteredthrough HPTS feed holes 174 into active convection cooling zone 173 andinto plenum 178 to facilitate impingement cooling in high pressureturbine section 112. Spent impingement cooling air 180 exits shroudsegment 130 through shroud segment leading edge cooling openings 181 tofacilitate purging gap 182 defined between high pressure turbine nozzleouter band 183 and shroud segment 130, and/or through cooling openings185 defined at a trailing end 205 of high pressure turbine section 112to facilitate film cooling inner surface 138 and/or 140 of shroudsegment 130.

Second portion 172 of high pressure cooling air 170 is directed intosecond active convection cooling zone 186 that is defined at leastpartially between shroud support 144 and shroud segment 130, and betweenmidsection position control ring 162 and aft position control ring 164.Second portion 172 of high pressure cooling air 170 facilitates coolinglow pressure turbine section 114. In one embodiment, second portion 172of high pressure cooling air 170 is metered through a plurality of lowpressure turbine feed holes 187 defined in shroud support 144. Morespecifically, second portion 172 of high pressure cooling air 170 ismetered directly into active convection cooling zone 186 to facilitateshroud segment impingement cooling in low pressure turbine section 114,such that cooling air bypasses a third region 210 defining an inactiveconvection cooling zone 211 between shroud support 144 and shroud ringstructure 158, and between midsection position control ring 162 and aftposition control ring 164. Spent impingement cooling air exits shroudsegment 130 through cooling openings 194 defined at or near trailingedge 137 of shroud segment 130.

In the flow path illustrated in FIG. 1, high pressure turbine sectionactive convection cooling zone 173 and/or low pressure turbine sectionactive convection cooling zone 186 are directly and actively cooled. Lowpressure turbine section inactive convection cooling zone 211 isinactive, i.e., no high pressure cooling air flows through inactiveconvection cooling zone 211. Thus, a thermal response within inactiveconvection cooling zone 211 to environmental conditions created duringtransient engine operation is reduced and/or retarded. As a result,transient displacement of midsection position control ring 162 and/oraft position control ring 164 is also reduced and/or retarded.

As shown in FIG. 2, high pressure cooling air 170 is directed into duct204 defined at least partially between high pressure turbine nozzleouter band 183 and shroud ring structure 158 forming midsection positioncontrol ring 162. High pressure cooling air 170 is separated into firstportion 171 and second portion 172. First portion 171 of high pressurecooling air 170 is metered through HPTS feed hole(s) 174 into highpressure turbine section active convection cooling zone 173 at leastpartially defining plenum 176 and plenum 178 to facilitate shroudsegment impingement cooling in high pressure turbine section 112. Spentimpingement cooling air 180 exits shroud segment 130 through shroudsegment leading edge cooling openings 181 to facilitate purging gap 182between high pressure turbine nozzle outer band 183 and shroud segment130 and/or through cooling openings 185 defined at trailing end 205 ofhigh pressure turbine section 112 to facilitate film cooling innersurface 138 and/or 140.

Second portion 172 of high pressure cooling air 170 is directed into lowpressure turbine section active convection cooling zone 186 defined atleast partially between shroud support 144 and shroud segment 130, andbetween midsection position control ring 162 and aft position controlring 164 to facilitate cooling low pressure turbine section 114. In oneembodiment, second portion 172 of high pressure cooling air 170 ismetered through a plurality of low pressure turbine feed holes 187defined through shroud support 144. Second portion 172 of high pressurecooling air 170 is metered directly into low pressure turbine sectionactive convection cooling zone 186 at least partially defining plenum189 and plenum 191 to facilitate shroud segment impingement cooling inlow pressure turbine section 114. Spent impingement cooling air 193exits shroud segment 130 through cooling openings 194 defined at or neartrailing edge 137 of shroud segment 130.

The shroud cooling assembly as shown in FIGS. 1 and 2 directs highpressure cooling air directly into high pressure turbine section activeconvection cooling zone 173 and/or low pressure turbine section activeconvection cooling zone 186 through respective feed hole(s) 174 and feedhole(s) 187.

In the shroud cooling assembly as shown in FIGS. 1 and 2, high pressurecooling air is not metered or directed through low pressure turbinesection inactive convection cooling zone 211. As a result, thecomponents defining low pressure turbine section inactive convectioncooling zone 211 respond relatively slower to thermal conditions and/orenvironments during transient engine operation than the componentsdefining an active convection cooling zone within conventional shroudcooling assemblies. This slower response to thermal conditions and/orenvironments facilitates relatively slower transient displacement ofmidsection position control ring 162 and/or aft position control ring164.

Thus, by bypassing the low pressure turbine section shroud ringstructure, the high pressure cooling air flow paths shown in FIGS. 1 and2 facilitate reducing and/or retarding the transient thermal responseand/or displacement of the shroud segment during transient engineoperation. The slower response further facilitates improved blade tipclearance and turbine engine efficiency.

FIG. 3 illustrates a portion of an exemplary turbine nozzle and shroudassembly 300 for use in a gas turbine engine. Shroud assembly 300 issimilar to shroud assembly 110, and as such, components illustrated inFIG. 3 that are identical to components illustrated in FIGS. 1 and 2 areidentified in FIG. 3 using the same reference number used in FIGS. 1 and2. Gap 182 is defined at an interface between outer band 183 of upstreamturbine nozzle 112 and an immediately downstream and adjacent shroudassembly 110 that includes a shroud segment 302. In the exemplaryembodiment, turbine nozzle 112 is coupled to shroud segment 302 to formturbine nozzle and shroud assembly 300.

A shroud segment leading edge 304 defines a forward face 306 of shroudsegment 302. In the exemplary embodiment, a radially inner surface 308of shroud segment 302 extends substantially perpendicularly to forwardface 306. Further, in the exemplary embodiment, a rounded or arcuatecorner portion 310 extends between forward face 306 and inner surface308 and partially defines gap 182. Moreover, in the exemplaryembodiment, a circumferential plenum 312 is defined within leading edge304 and is in flow communication with gap 182.

Outer band 183 has a trailing edge 314 that defines an aft face 316 ofouter band 183. When turbine nozzle 112 is coupled to shroud segment302, gap 182 is generally defined between aft face 316 and forward face306. Gap 182 enables cooling air to flow radially inwardly towards acombustion gas or hot gas flow path 320. Hot gas flow path 320 flowsgenerally parallel to a central axis 322 defined by the gas turbineengine. The cooling air may include leakage air 324 directed from duct204, shown in FIG. 1, and/or shroud leading edge cooling air 326discharged from active convection cooling zone 173, shown in FIG. 1.

In the exemplary embodiment, shroud segment leading edge 304 includes atleast one forwardly directed cooling opening 360 that extendssubstantially perpendicularly therethrough. Specifically, coolingopening 360 extends from plenum 312 through leading edge 304. Moreover,in the exemplary embodiment, at least one discrete convection coolinghole 362 extends between plenum 312 and inner surface 308. Specifically,in the exemplary embodiment, convection cooling hole 362 extendsobliquely from plenum 312 through leading edge 304 to inner surface 308at a location that is radially inward from cooling opening 360.

During operation, cooling opening 360 meters a flow of shroud leadingedge cooling air 326 into gap 182. Cooling air 326 is discharged fromcooling opening 360 into gap 182, wherein the air 326 mixes with leakageair 324 discharged from duct 204, shown in FIG. 1. Plenum 312facilitates circumferentially mixing leakage air 324 and air 326 toprovide a substantially uniform distribution of mixed cooling air 324throughout gap 182. A portion of the mixed cooling air 364 is channeledthrough the at least one discrete convection cooling hole 362 to innersurface 308 to facilitate forming a recuperated film cooling layer 366across inner surface 308. The remaining mixed cooling air 364 isdischarged from gap 182 to facilitate film cooling of corner portion310. Film cooling layer 366 facilitates forming a cooling barrierbetween hot gas flow path 320 and shroud segment 302. Moreover, coolingair 364 discharged from hole 362 provides energy to cooling layer 366such that film cooling layer 366 is facilitated to be maintained inproximity to surface 308 for a longer distance across surface 308 priorto separating from surface 308. As a result, an operating temperature ofshroud segment 302 is reduced such that a useful life of the turbineengine and/or shroud segment is facilitated to be increased.

FIG. 4 illustrates an alternative embodiment of turbine nozzle andshroud assembly 300. The alternative embodiment illustrated in FIG. 4 issubstantially similar to the embodiment shown in FIG. 3. As such,components illustrated in FIG. 4 that are identical to componentsillustrated in FIG. 3 are identified in FIG. 4 using the same referencenumber used in FIG. 3. Gap 182 is defined at an interface between anouter band 400 of an upstream turbine nozzle 402 and an immediatelydownstream and adjacent shroud assembly 110 that includes a shroudsegment 302. In the exemplary embodiment, turbine nozzle 402 is coupledto shroud segment 302 to form turbine nozzle and shroud assembly 300 fora gas turbine engine. In the exemplary embodiment, a lip 404 is formedon an aft face 406 of turbine nozzle 402. More specifically, in theexemplary embodiment, lip 404 is positioned radially inwardly from gap182 and corner portion 310, and extends downstream, generally parallelto central axis 322, from aft face 406.

In the exemplary embodiment, at least one discharge opening 408 isdefined in aft face 406. Specifically, discharge opening 408 extendsobliquely through aft face 406 and is radially outward from lip 404.Discharge opening 408 meters the flow of spent turbine nozzle coolingair 410 against lip 404. Specifically, spent turbine nozzle cooling air410 is metered towards lip 404 from a turbine nozzle active convectioncooling zone 412 that is at least partially defined by outer band 400.

During operation, spent turbine nozzle cooling air 410 is dischargedfrom discharge opening 408 in a metered flow directed towards lip 404.Lip 404 re-directs cooling air 410 towards corner portion 310 and innersurface 308. Moreover, cooling air 364, which has been uniformlydistributed by plenum 312, is channeled through convection cooling hole362 towards inner surface 308. Cooling air 364 and cooling air 410 mixadjacent inner surface 308 to facilitate forming a recuperated filmcooling layer 414 across inner surface 308. Further, because lip 404extends downstream from aft face 406, downstream from gap 182, lip 404enables gap 182 to be pressurized to facilitate film cooling layer 414being formed. Film cooling layer 414 facilitates forming a coolingbarrier between hot gas flow path 320 and shroud segment 302, such thatundesirable hot gas injection into gap 182 due to nozzle trailing edgewake effect is facilitated to be reduced. Moreover, film cooling layer414 facilitates reducing an operating temperature of shroud segment 302,such that a useful life of the turbine engine and/or shroud assembly 110is facilitated to be increased.

Referring to FIGS. 3 and 4, in the exemplary embodiments, coolingopening 360 extends through shroud segment leading edge 304 to meter aflow of shroud leading edge cooling air 326 into gap 182. Moreover, inthe exemplary embodiments, cooling opening 360 is radially outward fromconvection cooling hole 362 such that shroud leading edge cooling air326 discharged into gap 182 from cooling opening 360 is mixed withleakage air 324 directed from duct 204 (shown in FIG. 1). Plenum 312facilitates circumferentially mixing shroud leading edge cooling air 326and leakage air 324 to provide a substantially uniform distribution ofmixed cooling air 364 within gap 182. A portion of mixed cooling air 364is channeled through convection cooling hole 362 towards inner surface308 to facilitate forming a recuperated film cooling layer 366 acrossinner surface 308. In one embodiment, lip 404 is positioned radiallyinward from at least one discharge opening 408, such that spent turbinenozzle cooling air 410 discharged from opening 408, is re-directed bylip 404 across corner portion 310 and towards inner surface 308. Mixedcooling air 364 discharged from discrete convection cooling hole 362,and cooling air 410 discharged from opening 408, are mixed adjacentsurface 308 to facilitate forming a recuperated film cooling layer 414across inner surface 308. Film cooling layers 366 and 414 facilitateforming a cooling barrier between hot gas flow path 320 shroud segment302, such that undesirable hot gas injection into gap 182 due to nozzletrailing edge wake effect is facilitated to be reduced. Moreover, filmcooling layers 366 and 414 facilitate reducing an operating temperatureof shroud segment 302. As such, film cooling layers 366 and 414facilitate increasing a useful life of the turbine engine and/or shroudassembly 110.

The above-described methods and systems facilitate film cooling a shroudsegment. The methods and systems facilitate forming a barrier betweenthe hot gas flow path flowing through the high pressure turbine andcooling air flowing through and exiting a gap defined between theturbine nozzle and the shroud segment. More specifically, cooling airflowing through the gap is directed through an opening extending througha leading edge of the shroud segment to facilitate forming a coolinglayer on the shroud segment inner surface. In an alternative embodiment,spent turbine nozzle cooling air exiting the turbine nozzle outer bandis directed to impinge against a lip extending from the trailing edge ofthe turbine nozzle. The lip is positioned radially inward from the gapand extends downstream from the gap to direct the cooling air towards arounded corner portion formed on the leading edge of the shroud segment.The corner facilitates forming or developing a film cooling layer at,near, or adjacent to, the inner surface of the shroud segment downstreamof the gap. Specifically, the film layer is formed from a combination ofcooling air directed by the lip and corner portion and cooling airdischarged from the opening that extends through the shroud segmentleading edge. In one embodiment, the extended lip serves as a barrierbetween the hot gas flow path and the cooling air flowing through andexiting the gap defined between the outer band and the shroud segment.Further, the extended lip facilitates pressurizing the cooling airwithin the gap to substantially prevent or limit hot gas injection intothe gap due to nozzle trailing edge wake effect as may be seen inconventional cooling systems or assemblies. Moreover, the film coolinglayer facilitates reducing a temperature of the shroud segment, suchthat the film cooling layer facilitates increasing a life-span of theturbine engine and/or increasing an efficiency of the turbine engine.

In one embodiment, a method for cooling a shroud segment of a gasturbine engine is provided. The method includes providing a turbineshroud assembly including a shroud segment having an inner surface and aleading edge that is substantially perpendicular to the inner surface,and coupling a turbine nozzle to the turbine shroud segment such that agap is defined between an aft edge of an outer band of the turbinenozzle and the leading edge. The method also includes directing coolingair into the gap, circumferentially mixing the cooling air in a plenumdefined within the leading edge to substantially uniformly distributethe cooling air throughout the gap, and directing the cooling air in thegap through at least one cooling hole formed between the plenum and theinner surface.

Exemplary embodiments of methods and systems for film cooling a shroudsegment are described above in detail. The method and system are notlimited to the specific embodiments described herein, but rather, stepsof the method and/or components of the system may be utilizedindependently and separately from other steps and/or componentsdescribed herein. Further, the described method steps and/or systemcomponents can also be defined in, or used in combination with, othermethods and/or systems, and are not limited to practice with only themethod and system as described herein.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A method for cooling a shroud segment of a gas turbine engine, saidmethod comprising: providing a turbine shroud assembly including ashroud segment having an inner surface and a leading edge that issubstantially perpendicular to the inner surface; coupling a turbinenozzle to the turbine shroud segment such that a gap is defined betweenan aft edge of an outer band of the turbine nozzle and the leading edge;directing cooling air into the gap such that the cooling air issubstantially circumferentially mixed in a plenum defined within theleading edge; and directing the cooling air in the gap through at leastone cooling hole formed between the plenum and the inner surface.
 2. Amethod in accordance with claim 1 wherein directing the cooling air inthe gap through at least one cooling hole formed between the plenum andthe inner surface further comprises directing the cooling air toward theinner surface to facilitate forming a recuperated film cooling layeracross the inner surface.
 3. A method in accordance with claim 1 whereindirecting cooling air into the gap further comprises metering a flow ofcooling air into the gap through at least one forwardly directed coolingopening defined in the plenum and positioned radially outward from theat least one cooling hole extending between the plenum and the innersurface.
 4. A method in accordance with claim 1 wherein directingcooling air into the gap further comprises directing cooling air intothe gap such that the air is substantially uniformly distributedthroughout the gap from the plenum.
 5. A method in accordance with claim1 wherein a lip is formed on the aft edge and extends downstream fromthe aft edge, said directing cooling air into the gap further comprises:metering spent turbine nozzle cooling air into the gap through at leastone discharge opening extending through the aft edge; and directing thespent turbine nozzle cooling air in the gap against the lip and towardsthe shroud segment to facilitate cooling the shroud segment.
 6. A methodin accordance with claim 5 wherein the shroud segment includes a roundedcorner portion extending between the leading edge and the inner surface,said method further comprises directing the spent turbine nozzle coolingair along the rounded corner portion to facilitate forming a filmcooling layer across the inner surface.
 7. A method in accordance withclaim 1 further comprising pressurizing the gap to facilitate reducing aturbine nozzle wake effect within the gap.
 8. A turbine nozzle andshroud assembly for a gas turbine engine, said turbine nozzle and shroudassembly comprising: a shroud segment comprising a leading edge and aninner surface that is substantially perpendicular to said leading edge;and a turbine nozzle comprising an outer band comprising an aft edge,said turbine nozzle is upstream from said shroud segment and is coupledwith said shroud segment such that a gap is defined between said aftedge and said leading edge, said shroud segment comprises: acircumferential plenum defined within said leading edge, said plenum isconfigured to substantially uniformly distribute cooling aircircumferentially throughout the gap; and at least one cooling holeformed between said plenum and said inner surface, said at least onecooling hole is configured to direct the cooling air in said gap towardthe hot gas flow path.
 9. A turbine nozzle and shroud assembly inaccordance with claim 8 wherein said at least one cooling hole isfurther configured to direct cooling air in the gap towards said innersurface to facilitate forming a recuperated film cooling layer acrosssaid inner surface.
 10. A turbine nozzle and shroud assembly inaccordance with claim 8 further comprising at least one forwardlydirected cooling opening defined in said plenum, said at least oneforwardly directed cooling opening is radially outward from said atleast one cooling hole formed between said plenum and said inner surfaceand is configured to meter cooling air into said gap.
 11. A turbinenozzle and shroud assembly in accordance with claim 8 further comprisinga duct at least partially defined by said turbine nozzle and saidturbine shroud segment, said duct configured to direct cooling air intosaid gap.
 12. A turbine nozzle and shroud assembly in accordance withclaim 8 further comprising: a lip formed on said aft edge, said lipextends downstream from said gap; and at least one discharge openingextending through said aft edge, said at least one discharge openingconfigured to meter spent turbine nozzle cooling air into said gap, suchthat said lip directs the spent turbine nozzle cooling air from said gaptowards said shroud segment to facilitate cooling said shroud segment.13. A turbine nozzle and shroud assembly in accordance with claim 12further comprising a rounded corner portion extending between saidleading edge and said inner surface, said rounded corner portionconfigured to direct the spent turbine nozzle cooling air towards saidinner surface to facilitate forming a film cooling layer against saidinner surface.
 14. A turbine nozzle and shroud assembly in accordancewith claim 8 wherein said gap is pressurized to facilitate minimizing aturbine nozzle wake effect within said gap.
 15. A cooling system for agas turbine engine that includes a shroud segment having a leading edgeand an inner surface that is substantially perpendicular to the leadingedge, and a turbine nozzle that is upstream from the shroud segment andthat includes an outer band having an aft edge, said cooling systemcomprising: a plenum defined within the leading edge, said plenum isconfigured to substantially uniformly distribute cooling air through agap defined between the aft edge and the leading edge; and at least onecooling hole formed between said plenum and the inner surface, said atleast one cooling hole is configured to direct the cooling air from thegap towards a hot gas flow path defined within the gas turbine engine.16. A cooling system in accordance with claim 15 further configured todirect cooling air from the gap through said at least one cooling holeextending between said plenum and the inner surface to facilitateforming a recuperated film cooling layer across the inner surface of theshroud segment.
 17. A cooling system in accordance with claim 15 furthercomprising at least one forwardly directed cooling opening defined insaid plenum, said at least one forwardly directed cooling opening isradially outward from said at least one cooling hole extending betweensaid plenum and the inner surface, said at least one forwardly directedcooling opening is configured to meter cooling air into the gap.
 18. Acooling system in accordance with claim 15 further comprising a duct atleast partially defined by the turbine nozzle and the turbine shroudsegment, said duct configured to direct cooling air into the gap.
 19. Acooling system in accordance with claim 15 further comprising: a lipformed on the outer band aft edge, said lip extends downstream from thegap; and at least one discharge opening extending through the aft edge,said at least one discharge opening is configured to meter spent turbinenozzle cooling air into the gap, said lip is configured to direct thespent turbine nozzle cooling air to facilitate film cooling the shroudsegment.
 20. A cooling system in accordance with claim 19 furthercomprising a rounded corner portion extends between the leading edge andthe inner surface, said rounded corner portions configured to directspent turbine nozzle cooling air towards the inner surface of the shroudsegment to facilitate forming a film cooling layer across the innersurface.